Higher operating temperatures of gas turbine engines are continually being sought in order to increase the efficiency of the engines. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of nickel, cobalt and iron based superalloys. These superalloys can be designed to withstand temperatures in the range of about 1000 to about 1100° C. or higher. Nonetheless, when used to form components of the turbine, such as combustor liners, augmentor hardware, shrouds and high and low-pressure nozzles and blades, the superalloys alone could be susceptible to damage by oxidation and hot corrosion attack. Accordingly, these components are typically protected by an environmental and/or a thermal barrier coating (“TBC”). In general, TBCs can be used in conjunction with the superalloys in order to reduce the cooling air requirements associated with a given turbine. Ceramic materials, such as yttrium-stabilized zirconia (YSZ), are widely used as a TBC or topcoat of TBC systems. These materials are employed because, for example, they can be readily deposited by plasma-spraying and physical vapor deposition (PVD) techniques, and they also generally exhibit desirable thermal characteristics. In general, these TBCs can be utilized in conjunction with the superalloys in order to reduce the cooling air requirements associated with a given turbine.
In order to be effective, TBCs need to possess low thermal conductivity, strongly adhere to the component and remain adhered through many heating and cooling cycles. The latter requirement is particularly demanding due to the different coefficients of thermal expansion between the ceramic materials and the superalloy substrates that they protect. To promote adhesion and extend the service life of a TBC, an oxidation-resistant bond coating typically takes the form of a diffusion aluminide coating or an overlay coating, such as MCrAlX where M is iron, cobalt and/or nickel and X is yttrium or another rare earth element. During the deposition of a ceramic TBC and subsequent exposures to high temperatures, such as during engine operation, these bond coats form a tightly adherent alumina (Al2O3) layer or scale that adheres the TBC to the bond coat.
The service life of a TBC is typically limited by a spallation event brought on by, for example, thermal fatigue. Accordingly, a significant challenge has been to obtain a more adherent ceramic layer that is less susceptible to spalling when subjected to thermal cycling. Though significant advances have been made, there is the inevitable requirement to repair components whose thermal barrier coatings have spalled. Though spallation typically occurs in localized regions or patches, a conventional repair method has been to completely remove the TBC after removing the affected component from the turbine or other area, restore or repair the bond coat as necessary and recoat the engine component. Techniques for removing TBCs include grit blasting or chemically stripping with an alkaline solution at high temperatures and pressures. However, grit blasting is a slow, labor-intensive process and can erode the surface beneath the coating. The use of an alkaline solution to remove a TBC also is less than ideal because the process typically requires the use of an autoclave operating at high temperatures and pressures. Consequently, some conventional repair methods are labor intensive and expensive, and can be difficult to perform on components with complex geometries, such as airfoils and shrouds. As an alternative, U.S. Pat. No. 5,723,078 to Nagaraj et al. teach selectively repairing a spalled region of a TBC by texturing the exposed surface of the bond coat, and then depositing a ceramic material on the textured surface by plasma spraying. While avoiding the necessity to strip the entire TBC from a component, the repair method taught by Nagaraj et al. requires removal of the component in order to deposit the ceramic material.
In the case of large power generation turbines, completely halting power generation for an extended period of time in order to remove components whose TBCs have suffered only localized spallation is not economically desirable.
U.S. Pat. No. 7,476,703 discloses an in-situ method and composition for repairing a thermal barrier coating, which is based on a silicone resin system. U.S. Pat. No. 6,413,578 discloses an in-situ method for repairing thermal barrier coating with a ceramic paste. In situ methods of repairing a damaged component, such as TBC coating, are also disclosed in U.S. Pat. Nos. 7,509,735, 8,563,080, and U.S. Patent Application Publication No. 2015/0174837. A repair composition is disclosed in U.S. Pat. No. 6,875,464. A commercially available repair composition, AIM-MRO SR Resin Patch, may also be used for TBC repair.
However, there remains a need for an apparatus that would allow for effective application of repair composition to damaged regions of TBC. Such effective application includes application of a repair composition not only to flat surfaces but also to non-planar curved surfaces, which also suffer from TBC damage in turbine assemblies. Accordingly, the present invention seeks to provide a novel squeegee apparatus designed for dispensing and effectively applying a repair composition and methods of use thereof.